The present invention generally concerns a thermally isolated deployable meteoroid/debris shield for spacecraft in general, and more particularly concerns a weight efficient deployable shield of segments providing 360 degree coverage of a cylindrically shaped spacecraft such as Space Station, while at the same time providing a superior thermal barrier to conductive and radiative heat losses to "cold space" from the spacecraft's internal environment.
Spacecraft destined for long term orbital use must be provided with meteoroid/debris impingement protection. Several factors necessitate this protection: First, the statistical likelihood of meteoroid impingement with long term use; Second, the increasingly larger amounts of orbital space debris released from earth launched spacecraft activity; and Third, the general trend to enhance meteoroid/debris impingement survivability reliability from safety, cost, and weight standpoints.
Additionally, it is an important concern to provide a thermal barrier to internal environment heat loss to "cold space" through conductive and radiative modes, or interface points, existing between the pressurized environment and the meteoroid shield, whether the shield is deployed or fixed. Typically, this has not been as critical of a concern. However, with the advent of Space Station and the opportunity for sustained manned space missions, an effective thermal barrier is a critical requirement. This concern pertains to environmental control system (ECS) sizing, control, power, weight, etc., as well as dew formation, cleanliness, and microbial growth environments. Thus it is important to minimize the interface conductivity through reduction of conduction paths and/or lowering of interface structure/mechanism conductivity, for example through isolation and/or insulation of these interface points.
There are several examples of other methods providing a thermally isolated shielding protection scheme. However, these methods were neither designed for nor adequate for the stringent weight, thermal, and ballistic protection requirements that are so intensive to the thirty year mission of the space station. An example of one such scheme can be seen in the Skylab protection device. Skylab utilized eight individual rigid panels per circumference, or circular cross section. Each individual panel consisted of skin, structure, crank links, prime mover, as well as peripheral elements such as brackets, bearing blocks, light seals, closeouts, skirting, etc. The entire system was retained during ascent with highly loaded, pyro-released tension bands. The system was inherently "heavy" due to the aerodynamic loading and restraint scheme. The eight panel deployment scheme was somewhat complex and possessed a large quantity of components, resulting in reduced reliability of the system.
Another method of providing meteoroid protection is to size the pressurized environment wall thickness to withstand penetration. This method, being extremely heavy, is simply inadequate from a weight sufficiency standpoint for utilization with Space Station.
The use of fixed shields in general is known in the art, but this method is also inadequate. Fixed shields require a large quantity of fastening interfaces resulting in poor thermal characteristics. The fixed shields also exhibit poor ballistic impingement properties.
U.S. Pat. No. 4,314,682 to Barnett et al. discloses a spring loaded mechanism for deploying a shield from a space vehicle. The means for deploying the shield includes a plurality of elongated spring members extending outwardly from the body of the space vehicle and a plurality of curved ribs having their ends connected to ends of the extending elongated members. When deployed, the shield is in the general form of an open shell or bathtub-like structure with end caps at each end, the space vehicle residing within the structure. The shielding material itself is highly flexible radar attenuating material which, prior to deployment, is packed and folded into a jettisonable pod carried along one side of the space vehicle. When the shield is stowed, the spring members are wrapped downwardly around the circumference of the vehicle and held in place by the pod. The shield spontaneously deploys when the pod is jettisoned releasing the spring members which are attached to the shielding material.
U.S. Pat. No. 4,919,366 to Cormier discloses a heat resistive wall assembly for space vehicles comprising an inner wall of wrought beryllium or aluminum providing structural support for the vehicle, and an outer wall of interlocking panels of a honeycomb laminate of heat resistive material. An evacuated jacket of insulating material is disposed between the inner and outer walls. The space between the inner and outer walls that is not contained within the evacuated jacket is vented to ambient atmosphere.
U.S. Pat. No. 4,578,920 to Bush et al. discloses deployable truss structure having first and second spaced surface truss layers. A passive spring positioned about an elongated shaft serves as the expansion force to move the folded struts from a stowed collapsed position to a deployed operative final truss configuration.
U.S. Pat. No. 4,166,598 to Seifert et al. discloses a stowable and inflatable vehicle enshrouding apparatus adapted for use in retaining heat emitted by a large, relatively hot space vehicle. The apparatus includes an inflatable framework external of which is attached a multilayer superinsulating blanket shroud attached to the inflatable members which comprise the frame. The apparatus is deployed by removing it from stowage and inflating the inflatable support members.
U.S. Pat. No. 4,164,339 to McClenny discloses an environmental protection system comprising sheets of thermal insulators superposed one upon the other and deployed over the surface to be protected. A "dead space" thermal insulation, such as a vacuum or simply a high resistance physical separation, is provided between the surfaces. The separation is effected by flaps actuated by an aerodynamic or forced airstream or by the static energy stored in uncoiled flaps. The insulating material is stored on reels and deployed by automatic or manual means.
U.S, Patent No. 4,009,851 to Cable discloses a spacecraft structure having a hollow inner cylindrical member and a plurality of planar bulkheads secured to the outer surface of the inner member and extending radially outward from the inner member. A plurality of planar enclosure panels are secured to the extended edges of the bulkheads and each other to form and enclosed spacecraft structure.
U.S. Pat. No. 4,730,797 to Minovitch, U.S. Pat. No. 3,547,375 to Mackey, and U.S. Pat. No. 3,064,317 to Dobson all relate generally to the field of the present invention.